Gas turbine engine with gearbox health features

ABSTRACT

A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan and a braking system. The braking system is configured to selectively engage the fan during ground windmilling to apply a first level of braking to slow rotation of the fan. Further, when the rotation of the fan sufficiently slows, the braking system is further configured to apply a second level of braking more restrictive than the first level of braking.

BACKGROUND

This disclosure relates to health features for gas turbine engineshaving a geared architecture.

A gas turbine engine typically includes a fan section, a compressorsection, a combustor section, and a turbine section. Air entering thecompressor section is compressed and delivered into the combustorsection where it is mixed with fuel and ignited to generate a high-speedexhaust gas flow. The high-speed exhaust gas flow expands through theturbine section to drive the compressor and the fan section.

One type of gas turbine engine includes a geared architecture used todecrease the rotational speed of the fan. In one configuration, thegeared architecture includes a plurality of gears contained within agearbox. The gearbox is supplied with a flow of lubricating fluid, whichis typically oil, to protect the gears during operation.

The geared architecture not only must function during aircraftoperation. There are also challenges when the gas turbine engine isunlit (i.e., turned off) and the aircraft is stationary on the ground.Rotation of the unlit engine, known as windmilling, occurs as air flowsthrough the unlit engine and causes rotation of a portion of the engine.

To prevent windmilling when an aircraft is on-ground (referred to hereinas “ground windmilling”), caps are sometimes provided over the inlet andoutlet of the engine. The caps serve to prevent air flow, such as wind,from entering the unlit engine and causing windmilling. Other knownmethods of preventing windmilling include using a wedging device intothe gas turbine engine. Another known method uses a generator associatedwith the gas turbine engine as a dynamic brake, and attempts to use theresistance of the generator to slow the rotation of the engine.

SUMMARY

A gas turbine engine according to an exemplary aspect of the presentdisclosure includes, among other things, a fan and a braking system. Thebraking system is configured to selectively engage the fan during groundwindmilling to apply a first level of braking to slow rotation of thefan. Further, when the rotation of the fan sufficiently slows, thebraking system is further configured to apply a second level of brakingmore restrictive than the first level of braking.

In a further non-limiting embodiment of the foregoing engine, the engineincludes a gear reduction between the fan and a shaft of the engine.

In a further non-limiting embodiment of the foregoing engine, thebraking system includes a brake, and the fan includes a disc. The brakeis configured to selectively engage the disc to apply the first level ofbraking to slow rotation of the fan.

In a further non-limiting embodiment of the foregoing engine, thebraking system further includes a pawl, and the disc includes a slot.The pawl is configured to selectively engage the slot to apply thesecond level of braking and to substantially lock the fan againstrotation.

In a further non-limiting embodiment of the foregoing engine, the brakeengages the disc following an engine shut-off event and slows therotation of the fan. Further, when the rotation of the fan sufficientlyslows, the pawl engages the slot to lock the fan against furtherrotation.

In a further non-limiting embodiment of the foregoing engine, the pawlis biased into engagement with the slot, and the pawl is selectivelyretracted from the slot by an actuator.

In a further non-limiting embodiment of the foregoing engine, the pawland brake are designed to fail if deployed during an in-flightcondition.

Another gas turbine engine according to an exemplary aspect of thepresent disclosure includes, among other things, a fan and a lubricationsystem. The lubrication system is configured to pump lubricant into afan drive gearbox when the fan is windmilling at any rotational speedand direction.

In a further non-limiting embodiment of the foregoing engine, thelubrication system is configured to pump lubricant to the fan drivegearbox when the fan rotates below 1,000 rpm.

In a further non-limiting embodiment of the foregoing engine, thelubrication system includes a main pump and a main reservoir fluidlycoupled to the main pump. The main pump is configured to pump lubricantfrom the main reservoir to the fan drive gearbox during normal operatingconditions. Further, the lubrication system includes a secondary pumpand a secondary reservoir fluidly coupled to the secondary pump. Thesecondary pump is configured to pump lubricant from the secondaryreservoir to the fan drive gearbox when the main pump does not provideadequate lubricant to the fan drive gearbox.

In a further non-limiting embodiment of the foregoing engine, thesecondary pump is one of (1) an accessory pump, (2) a rotary-shaftdriven pump, (3) an electrically-driven pump, and (4) an aircrafthydraulic system-powered pump.

A further gas turbine engine according to an exemplary aspect of thepresent disclosure includes, among other things, a fan, a gearedarchitecture coupled to the fan, at least one sensor configured togenerate signals indicative of a condition of the geared architecture,and a control unit electrically coupled to the at least one sensor. Thecontrol unit is configured to interpret signals from the at least onesensor to determine a condition of the geared architecture.

In a further non-limiting embodiment of the foregoing engine, the engineincludes a lubrication system configured to pump lubricant into agearbox of the geared architecture when the fan is windmilling. The atleast one sensor is configured to generate signals indicative of acondition of the lubrication system.

In a further non-limiting embodiment of the foregoing engine, the atleast one sensor is a temperature sensor and is configured to generatesignals indicative of the temperature of the lubricant within thelubrication system.

In a further non-limiting embodiment of the foregoing engine, thetemperature sensor is adjacent a scavenge line of the lubricationsystem.

In a further non-limiting embodiment of the foregoing engine, the atleast one sensor is a pressure sensor and is configured to generatesignals indicative of the pressure of the lubricant within thelubrication system, and the pressure sensor is adjacent a supply line ofthe lubrication system.

In a further non-limiting embodiment of the foregoing engine, the atleast one sensor is a debris sensor configured to generate signalsindicative of a level of debris within the lubrication system, and thedebris sensor is adjacent a scavenge line of the lubrication system.

In a further non-limiting embodiment of the foregoing engine, the atleast one sensor is a vibration sensor configured to generate signalsindicative of a level of vibration of the geared architecture, and thevibration sensor is positioned adjacent a bearing near the gearedarchitecture.

Yet another gas turbine engine according to an exemplary aspect of thepresent disclosure includes, among other things, a fan, a gear reductionbetween the fan and a shaft of the engine, and a braking systemconfigured to selectively engage the fan to substantially preventwindmilling. The braking system is held in engagement with the fanwithout requiring electrical power.

In a further non-limiting embodiment of the foregoing engine, thebraking system includes a pawl, and the pawl is mechanically biased intoengagement with the fan by way of a spring.

The embodiments, examples and alternatives of the preceding paragraphs,the claims, or the following description and drawings, including any oftheir various aspects or respective individual features, may be takenindependently or in any combination. Features described in connectionwith one embodiment are applicable to all embodiments, unless suchfeatures are incompatible.

BRIEF DESCRIPTION OF THE DRAWINGS

The drawings can be briefly described as follows:

FIG. 1 schematically illustrates a gas turbine engine.

FIG. 2 schematically illustrates a braking system.

FIG. 3 is a flow chart representing an example method.

FIG. 4A is a graphical representation of a response of a gas turbineengine using the disclosed braking system.

FIG. 4B is a graphical representation of a response of a gas turbineengine without the disclosed braking system.

FIG. 5 schematically illustrates a gas turbine engine, and in particularillustrates a lubrication system and a plurality of sensors.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. Alternative engines mightinclude an augmentor section (not shown) among other systems orfeatures. The fan section 22 drives air along a bypass flow path B in abypass duct defined within a nacelle 15, while the compressor section 24drives air along a core airflow path C for compression and communicationinto the combustor section 26 then expansion through the turbine section28. Although depicted as a two-spool turbofan gas turbine engine in thedisclosed non-limiting embodiment, it should be understood that theconcepts described herein are not limited to use with two-spoolturbofans as the teachings may be applied to other types of turbineengines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The gearedarchitecture 48 in this example includes a gearbox G, which is ahousing, and encloses one or more gears, such as a sun gear and one ormore planetary gears. The high speed spool 32 includes an outer shaft 50that interconnects a second (or high) pressure compressor 52 and asecond (or high) pressure turbine 54. A combustor 56 is arranged inexemplary gas turbine 20 between the high pressure compressor 52 and thehigh pressure turbine 54. A mid-turbine frame 57 of the engine staticstructure 36 is arranged generally between the high pressure turbine 54and the low pressure turbine 46. The mid-turbine frame 57 furthersupports bearing systems 38 in the turbine section 28. The inner shaft40 and the outer shaft 50 are concentric and rotate via bearing systems38 about the engine central longitudinal axis A which is collinear withtheir longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive geared architecture 48 may be varied. Forexample, geared architecture 48 may be located aft of combustor section26 or even aft of turbine section 28, and fan section 22 may bepositioned forward or aft of the location of geared architecture 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft, with the engine at its best fuel consumption—also known as“bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is theindustry standard parameter of lbm of fuel being burned divided by lbfof thrust the engine produces at that minimum point. “Low fan pressureratio” is the pressure ratio across the fan blade alone, without a FanExit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosedherein according to one non-limiting embodiment is less than about 1.45.“Low corrected fan tip speed” is the actual fan tip speed in ft/secdivided by an industry standard temperature correction of [(Tram°R)/(518.7°R)]^(0.5). The “Low corrected fan tip speed” as disclosedherein according to one non-limiting embodiment is less than about 1150feet/second (350.5 meters/second).

FIG. 2 schematically illustrates an example braking system 60 for usewith a gas turbine engine, such as the gas turbine engine 20. Thebraking system 60 is configured to selectively engage the fan 42 duringground windmilling to slow rotation of the fan 42 and, once sufficientlyslowed, to lock the fan 42 against rotation. While reference is made tothe gas turbine engine 20 of FIG. 1, the braking system 60 could be usedwith other gas turbine engines.

In this example, the braking system 60 includes a brake 62 configured toselectively engage a disc 64 of the fan 42. The disc 64 may beintegrally formed with the hub of the fan 42, or, alternatively,attached to a hub of the fan 42. The brake 62, in one example, is aspring-loaded disc-brake and includes at least one brake pad 66configured to selectively engage a face 68 of the disc 64. In anotherexample, the brake 62 includes two opposed pads configured to engagerespective fore and aft faces of the disc 64.

The brake 62 includes an actuator 70, shown schematically, coupled tothe brake pad 66. The actuator 70 is configured to adjust a position ofthe brake pad 66 to selectively engage the disc 64, per instructionsfrom a control unit U. As will be discussed below, the brake 62 isconfigured to apply a first level of braking to the disc to slowrotation of the fan 42 during ground windmilling.

The control unit U may be any known type of control unit, includingmemory, hardware, and software. In some examples, the control unit U maybe a stand-alone control unit, and may receive instructions from anelectronic engine controller (EEC) or a full authority digital enginecontroller (FADEC). Alternatively, the control unit U could be providedby the EEC or FADEC themselves.

The braking system 60 further includes a lock 72 including a pawl 74configured to selectively engage a corresponding slot 76 in the disc 64to apply a second level of braking to the disc, which is morerestrictive than the first level of braking applied by the brake 62. Inone example, the second level of braking substantially locks the fan 42against rotation. In this example, the pawl 74 is positioned radiallyoutward (e.g., in the radial direction R, which is normal to the enginecentral longitudinal axis A) of the disc 64, and the slot 76 is providedin the radially outer face of the disc 64.

The pawl 74, in one example, is mechanically biased, by way of a spring78, toward the disc 64 and, when aligned, into engagement with the slot76. In this example, the pawl 74 is moveable out of engagement with thedisc 64 by way of an actuator 80, which receives instructions form thecontrol unit U. Alternatively, the pawl 74 does not need to bemechanically biased, and could be moveable based on the movement of theactuator 80 alone. By mechanically biasing the pawl 74, however, thereis an added benefit of not requiring electrical power to maintain theposition of the pawl 74 within the slot 76. This feature becomesparticularly useful when the aircraft is grounded for extended periodsof time, such as weeks or even months. The lock 72 is shown somewhatschematically.

In FIG. 2, only one brake 62 is shown. It should be understood thatadditional brakes can be provided about the disc 64. Likewise, whileonly one lock 72 is shown, additional locks can be provided about thedisc 64. Further, in FIG. 2, the disc 64 includes only three slots 76.The disc 64 is not limited to three slots 76, and could include anynumber of slots 76.

One example method of using the braking system 60 will now be describedrelative to FIG. 3 and FIGS. 4A-4B. FIG. 3 is a flow chart thatrepresents an example method 82 of using the braking system 60. FIGS.4A-4B are graphical representations of fan rpm (rotations per minute)versus time for both an engine including the disclosed braking system 60(FIG. 4A) and an engine without the disclosed braking system (FIG. 4B).

In the method 82, the control unit U first determines, at 84, whether anaircraft having the braking system 60 is on ground and exposed to low tohigh winds that may cause windmilling. In another example, the controlunit U will lock the rotor as part of each and every engine shutdown.After shutdown, the engine is considered “off” and is electricallydormant. In some examples, the control unit U commands the lock 72 tomove into engagement with the fan 42, but the lock 72 remains in placewithout requiring electrical power. The control unit U may deploy thelock 72 partially in response to a signal from a “squat switch” of anaircraft, which senses the compression of the landing gear. The controlunit U may further deploy the lock following a command to shut down theengine.

If the aircraft is not on ground and exposed to a headwind (i.e., theaircraft is in flight and the engine is either providing thrust or is“off”), the control unit U commands the actuators 70, 80 to bring thebrake 62 and pawl 74 out of engagement with the disc 64, at 86. On theother hand, if the aircraft is on ground and the engine is commanded to“off,” the control unit U, at 88, first commands the actuator 70 tobring the brake pad 66 into engagement with the disc 64. Doing so willslow the rotation of the disc 64. Once the disc 64 substantially slowsto a near-stop or a complete stop, the control unit U commands theactuator 80 to deploy the pawl 74 into engagement with the slot 76 tostop any further rotation of the disc 64.

Step 88 may be further appreciated with reference to FIG. 4A. FIG. 4Aincludes a line 90 representative of rotational speed (in rpm, asdetermined by the EEC, for example) of the fan 42 relative to time. Inthe example of FIG. 4A, an aircraft having a gas turbine engine equippedwith the braking system 60 reaches ground idle at point 92 after theaircraft has reached a gate following a flight, for example. At point94, the fuel is cut off to the engine, at which point the fan 42 beginswindmilling. At point 96, the control unit U instructs the actuator 70to cause the brake pad 66 to engage the disc 64 to assist in slowingrotation of the fan 42. At point 98, as the rotational speed of the fan42 slows to a near-stop or a complete stop, the control unit U instructsthe actuator 80 to cause the pawl 74 to engage the slot 76 to lock thedisc 64 (and, in turn, the fan 42) against further rotation.

Without the braking system 60, the fan 42 would begin to windmill ineither a forward direction, illustrated at line segment 100 in FIG. 4B,or a rearward direction (in the case of a tail wind), illustrated atline segment 102 in FIG. 4B. Prolonged ground windmilling can increasethe wear on the gears, ball bearings, bearing races, and all otherhardware associated with the geared architecture 48 and more generallythe inner shaft 40 (often called the “low spool”). The braking system 60thus extends the life of all of this hardware and is particularlybeneficial in the rare instances where an aircraft is parked for a longperiod of time, such as weeks or even months.

With reference back to FIG. 3, the method 82 includes a plurality ofsteps configured to manage the use of the braking system when theaircraft is in-flight. At 104, the control unit U determines whether theaircraft is in-flight. If so, the control unit U next determines, at106, whether the brake 62 or the lock 72 have failed. A failure could beindicated by an unexpected position indications from a respectiveactuator 70, 80.

During normal in-flight operation of an aircraft, the brake 62 and lock72 are retracted, at 108 (i.e., if there is no failure of the brake orlock 62, 72). If a failure of the lock 72 exists such that the lock 72becomes engaged with the disc 64 during flight, for example, the lock 72is designed to fail in a manner to allow the rotor to turn. In oneexample, the pawl 74 is sized and/or made of a material selected toimmediately shear away and depart the assembly in a benign manner, at110, if engaged with the disc 64 during flight. If such a failurecondition occurs, a maintenance “flag” is delivered to the maintenancecomputer of the aircraft for follow up action by maintenance personnel.Alternatively, or additionally, if the brake 62 fails during flight andengages the disc 64, the brake 62 is also design to fail or simply wearaway. In one example, the brake pad 66 is sized and/or made of amaterial selected to be worn down relatively rapidly by the rotation ofthe disc 64 so as to not encumber operation of the fan 42, and, at 110,another maintenance “flag” is delivered to the pilot.

FIG. 5 schematically illustrates the gas turbine engine 20 and acorresponding lubrication system 112. In this example, the lubricationsystem 112 is configured to pump lubricant, such as oil, into thegearbox G (or, housing) containing one or more gears associated with thegeared architecture 48. In particular, the lubrication system 112 isconfigured to pump lubricant into the gearbox G when the gas turbineengine 20 is windmilling at any rotational speed and any direction. Inone particular example, the lubrication system 112 pumps lubricant tothe gearbox G when the gas turbine engine 20 is windmilling duringengine start before a main supply of oil typically pressurizes.

In this example, the lubrication system 112 includes a main pump 114 anda main reservoir 116 fluidly coupled to the main pump 114. The main pump114 is configured to pump lubricant from the main reservoir 116 into thegearbox G via a main supply line 118 during normal operating conditions.The term “normal operating conditions,” for purposes of this disclosure,includes conditions where the main pump provides an adequate amount oflubricant to the gearbox G. Lubricant is returned from the gearbox G tothe main reservoir 116 via a main scavenge (or, return) line 120. Themain reservoir 116 may be the main oil tank of the gas turbine engine20.

The lubrication system 112 further includes a secondary pump 122configured to pump lubricant to the gearbox G when the main pump 114does not provide adequate lubricant to the gearbox G. In this example,the secondary pump 122 is fluidly coupled to a secondary reservoir 124,and pumps lubricant to the gearbox G via a secondary supply line 126.Downstream of the gearbox G, lubricant is returned to the secondaryreservoir 124 via a secondary scavenge (or, return) line 128. In thisexample, the secondary supply and scavenge lines 126, 128 are separatefrom the main supply and scavenge lines 118, 120 in case there is anissue with the main supply and scavenge lines 118, 120. The secondaryreservoir 124 may be provided with lubricant from a main oil tank of thegas turbine engine 20, and may further be configured to isolate itselfautomatically from the main oil tank of the gas turbine engine if a lackof oil is detected.

In this example, the main pump 114 is a pump dedicated to providinglubricant to the gearbox G. The secondary pump 122, on the other hand,may be a pump that has a primary function other than to providelubricant to the gearbox G. In particular, the secondary pump 122 may beone of (1) an accessory pump, (2) a rotary-shaft driven pump, (3) anelectrically-driven pump, and (4) an aircraft hydraulic system-poweredpump.

In some examples, the main pump 114 may not be active during windmillingconditions. In those examples, the control unit U is configured toactivate the secondary pump 122 to provide a flow of lubricant into thegearbox G to protect the geared architecture 48 from wearing by movingwithout adequate lubricant.

In still other examples, a gas turbine engine may be programmed to beginproviding lubricant to the gearbox G via a main pump only when the fanrotates above a minimum threshold speed, such as 1,000 rpm. If such isthe case, the secondary pump 122 will be activated to provide lubricantto the gearbox at speeds below 1,000 rpm, for example.

In FIG. 5, there are a plurality of sensors associated with thelubrication system 112 and the gas turbine engine 20. The sensors areconfigured to generate signals indicative of a condition of the gearedarchitecture 48. Those signals are interpreted by the control unit U,which is configured to take an appropriate action in response to thereceived signals.

In FIG. 5, three sensors 130, 132, and 134 detect conditions associatedwith the lubrication system 112 (e.g., properties of the lubricant)related to the health of the geared architecture 48. A first sensor 130is a temperature sensor 130, and is configured to generate signalsindicative of the temperature of the lubricant within the lubricationsystem 112. In this example, the temperature sensor 130 is providedadjacent the scavenge line 120. The signals from the temperature sensor130 are interpreted by the control unit U to determine whether thegeared architecture 48 is being adequately cooled by the lubricationsystem 112. The temperature sensor 130 may be a thermistor or athermocouple, as examples.

A second sensor 132 is a pressure sensor, and is configured to generatesignals indicative of the pressure of the lubricant within thelubrication system 112. The pressure sensor 132, in this example, isadjacent a supply line of the lubrication system 112. The control unit Uis configured to interpret the signals from the pressure sensor 132 todetermine whether the main pump 112, for example, is adequatelypressurizing the lubricant flowing to the gearbox G. The pressure sensor132 may be provided by a pressure transducer, as one example.

A third sensor 134 is a debris sensor, and is configured to generatesignals indicative of a level of debris within the lubrication system.In this example, the debris sensor 134 is provided adjacent the scavengeline 120 to detect for the presence of debris that has been potentiallyexpelled from the gearbox G. For example, the debris sensor 134 maydetect the presence of metallic particles (e.g., copper, lead, orsilver) within the scavenge line 120, which may suggest that one or moregears within the gearbox G is being worn. In this respect, the debrissensor 134 may be provided by any type of sensor configured to detectthe presence of metal within a fluid.

In the example of FIG. 5, the gas turbine engine 20 includes a fourthsensor 136, which is a vibration sensor configured to generate signalsindicative of a level of vibration of the geared architecture 48. Inthis example, the vibration sensor 136 is positioned adjacent a bearing138 near the geared architecture 48. Increases in vibration at thebearing 138 can suggest that the health of the geared architecture 48 isdeclining. While the bearing 138 is illustrated generally, the vibrationsensor 136 could be positioned adjacent the #1 bearing of the engine 20,however the vibration sensor 136 is not limited to any particularlocation. The vibration sensor 136 can be provided by an accelerometer,for example.

In FIG. 5, only one of each type of sensor 130, 132, 134, and 136 isillustrated. It should be understood that additional sensors could beincluded. Further, while the sensors 130, 132, and 134 are shownrelative to the main supply line 118 and main scavenge line 120, similarsensors would also be provided relative to the secondary supply line 126and the secondary scavenge line 128 to detect for the same conditions asdescribed above. When adverse conditions are detected relative to themain supply and scavenge lines 118, 120, the control unit U may activatethe secondary pump 122 (if available) or a maintenance “flag” may bedelivered to the pilot. Further, while illustrated separately forpurposes of discussion herein, the braking system 60, lubrication system112 and the associated sensors 130, 132, 134, and 136 could be used incombination in a single engine 20.

It should be understood that terms such as “fore,” “aft,” “axial,”“radial,” and “circumferential” are used above with reference to thenormal operational attitude of the engine 20. Further, these terms havebeen used herein for purposes of explanation, and should not beconsidered otherwise limiting. Terms such as “generally,”“substantially,” and “about” are not intended to be boundaryless terms,and should be interpreted consistent with the way one skilled in the artwould interpret those terms.

Although the different examples have the specific components shown inthe illustrations, embodiments of this disclosure are not limited tothose particular combinations. It is possible to use some of thecomponents or features from one of the examples in combination withfeatures or components from another one of the examples.

One of ordinary skill in this art would understand that theabove-described embodiments are exemplary and non-limiting. That is,modifications of this disclosure would come within the scope of theclaims. Accordingly, the following claims should be studied to determinetheir true scope and content.

1. A gas turbine engine, comprising: a fan; and a lubrication systemconfigured to pump lubricant into a fan drive gearbox when the fan iswindmilling at any rotational speed and direction.
 2. The gas turbineengine as recited in claim 1, wherein the lubrication system isconfigured to pump lubricant to the fan drive gearbox when the fanrotates below 1,000 rpm.
 3. The gas turbine engine as recited in claim1, wherein: the lubrication system includes a main pump and a mainreservoir fluidly coupled to the main pump, the main pump configured topump lubricant from the main reservoir to the fan drive gearbox duringnormal operating conditions; and the lubrication system further includesa secondary pump and a secondary reservoir fluidly coupled to thesecondary pump, the secondary pump configured to pump lubricant from thesecondary reservoir to the fan drive gearbox when the main pump does notprovide adequate lubricant to the fan drive gearbox.
 4. The gas turbineengine as recited in claim 3, wherein the secondary pump is one of (1)an accessory pump, (2) a rotary-shaft driven pump, (3) anelectrically-driven pump, and (4) an aircraft hydraulic system-poweredpump.
 5. A gas turbine engine, comprising: a fan; a geared architecturecoupled to the fan; at least one sensor configured to generate signalsindicative of a condition of the geared architecture; and a control unitelectrically coupled to the at least one sensor, the control unitconfigured to interpret signals from the at least one sensor todetermine a condition of the geared architecture.
 6. The gas turbineengine as recited in claim 5, further comprising: a lubrication systemconfigured to pump lubricant into a gearbox of the geared architecturewhen the fan is windmilling, wherein the at least one sensor isconfigured to generate signals indicative of a condition of thelubrication system.
 7. The gas turbine engine as recited in claim 6,wherein the at least one sensor is a temperature sensor and isconfigured to generate signals indicative of the temperature of thelubricant within the lubrication system.
 8. The gas turbine engine asrecited in claim 7, wherein the temperature sensor is adjacent ascavenge line of the lubrication system.
 9. The gas turbine engine asrecited in claim 6, wherein: the at least one sensor is a pressuresensor and is configured to generate signals indicative of the pressureof the lubricant within the lubrication system; and the pressure sensoris adjacent a supply line of the lubrication system.
 10. The gas turbineengine as recited in claim 6, wherein: the at least one sensor is adebris sensor configured to generate signals indicative of a level ofdebris within the lubrication system; and the debris sensor is adjacenta scavenge line of the lubrication system.
 11. The gas turbine engine asrecited in claim 5, wherein: the at least one sensor is a vibrationsensor configured to generate signals indicative of a level of vibrationof the geared architecture; and the vibration sensor is positionedadjacent a bearing near the geared architecture.
 12. A gas turbineengine, comprising: a fan; a gear reduction between the fan and a shaftof the engine; and a braking system configured to selectively engage thefan to substantially prevent windmilling, and wherein the braking systemis held in engagement with the fan without requiring electrical power.13. The gas turbine engine as recited in claim 12, wherein the brakingsystem includes a pawl, the pawl mechanically biased into engagementwith the fan by way of a spring.